As is well known in the gas turbine engine technology, the efficiency of the engine is greatly enhanced by increasing the temperature of the turbine and/or reducing the amount of air that is required to maintain the turbine components within their tolerable limits. In other words, the material used for the turbine blades must be able to withstand the temperature and hostile environment that is seen in the turbine section. Engineers and scientist have been working for many years at improvements to provide materials capable of withstanding higher temperatures and to reduce the amount of coolant for achieving satisfactory cooling of the turbine components and particularly the turbine blade.
An example of cooled turbine blades is exemplified in U.S. Pat. No. 5,720,431 granted to Sellers, et al on Feb. 24, 1998 entitled COOLED BLADES FOR A GAS TURBINE ENGINE which teaches the use of feed chambers and feed channels where the feed channels extend from the root of the blade to the tip and include a discharge opening at the tip, the feed chamber connects to the source of coolant and through radial spaced impingement cooling holes replenishes the air in the feed channels. This teachings relate to the leading edge, trailing edge and the mid chord section. It is noted that this invention is principally concerned with the suction surface and the pressure surface in the mid chord section. This reference is incorporated herein by reference and should be referred to for a detailed description of air cooled turbine blades utilized in gas turbine engines.
U.S. Pat. No. 6,129,515 granted to Soechting, et al on Oct. 10, 2000 entitled TURBINE AIRFOIL SUCTION AIDED FILM COOLING MEANS is also included herein because not only does it describe cooled turbine blades, but it is particularly directed to teachings that is directed to means for slowing the velocity of the discharge air from the air film cooling holes so as to better disperse the air as it leaves the discharge ports and hence, tend to more effectively provide a film of cooling air adjacent to the outer surface at the pressure surface of the blade. It will be noted, for example, that the teaching includes step diffuser to attain a wider diffusion angle of the discharging film. This patent is also incorporated herein by reference.
U.S. Pat. No. 5,486,093 granted to Auxier et al on Jan. 23, 1996 entitled LEADING EDGE COOLING OF TURBINE AIRFOILS is included herein because it teaches the use of helix shaped cooing passages to enhance convective efficiency of the cooling air and to improve discharge of the film cooling air by orienting the discharge angle so that the discharging air is delivered more closely to the pressure and suction surfaces. The helix holes place the coolant closer to the outer surface of the blade to more effectively reduce the average conductive length of the passage so as to improve the convective efficiency. Also higher heat transfer coefficients are produced on the outer diameter of helix holes improving the capacity of the heat sink. This patent is likewise incorporated herein by reference.
As one skilled in this art will appreciate the heretofore design of cooled turbine blades typically utilize radial flow channels plus re-supply holes in conjunction with film discharge cooling holes as is exemplified in U.S. Pat. No. 5,720,431, supra. While this patent discloses a near wall cooling technique, this cooling construction approach has its downside because the hot gas temperature and pressure variation of the engine's working medium makes the control of the radial and chord-wise cooling flow difficult to achieve. A single pass radial channel flow as taught by the U.S. Pat. No. 5,720,431, supra, is not the ideal method of utilizing cooling air and as a consequence, this method results in a low convective cooling effectiveness.
The present invention obviates the problem noted in the above paragraph. The design philosophy of this invention as compared to the teachings noted above and the results obtained by the utilization of this invention as a cooling technique for turbine blades will enhance the cooling effectiveness and hence, will improve the efficiency of the engine. Essentially, this invention relates to cooling the surfaces of the pressure side and suction side of the airfoil and provides a matrix of square or rectangularly shaped cells (although other shapes could also be employed), each of which have discrete cooling passage(s) formed in the wall of the airfoil adjacent to the pressure surface and to the suction surface of the blade resulting in a near wall cooling technique of the turbine airfoil. The matrix can be made to span the longitudinal and chord-wise directions to encompass the entire pressure and suction surfaces or to a lesser portion depending on the heat load of a particular engine application. These cells not only can be arranged in an online array along the airfoil main body, the cells can also be a staggered array along the airfoil main body.
In addition, this invention contemplates the use of means for generating vortices in each of the passages to enhance heat transfer and the conductive characteristics of the cooling system. The multi-vortex cell of this invention serves to generate a high coolant flow turbulence level and, hence, yields a very high internal convection cooling effectiveness in comparison to the single pass construction described in the U.S. Pat. No. 5,720,431, supra.
In accordance with this invention, the designer can design each individual cell based on airfoil gas side pressure distribution in both the chord-wise and radial directions. Additionally each cell can be designed to accommodate the local external heat load on the airfoil so as to achieve a desired local metal temperature.
The discharge angle of the discharge passage of the vortex cooling passage is oriented to provide a film cooling hole where the discharge angle will enhance the film cooling effectiveness of the coolant. As will be appreciated by those familiar with this technology, film cooling on the suction side downstream of the gage point, i.e., the point where the two adjacent blades define the throat of the passage between blades, adversely affects the aerodynamics of film mixing and hence is a deficit in performance. This then becomes a trade-off in design to either obtain the benefits of film cooling in deference to these aerodynamic losses. To avoid the aerodynamic losses in heretofore known cooling schemes, in accordance with this invention cooling the suction side of the blade downstream of the gage point is provided by the airfoil internal multi-pass serpentine passage. This invention has the advantage over these schemes and hence is a significant improvement because the aft portion of the suction side wall of the airfoil can be internally cooled with the multi-vortex cell of this invention before discharging the coolant through the film discharge holes as a film upstream of the gage point in contrast to being discharged downstream of the gage point and thus, avoiding the aerodynamic losses associated with film mixing.